The invention described herein may be manufactured and used by or for the Government of the United States for all governmental purposes without the payment of any royalty.
The present invention relates generally to a method of repairing cracks in structures and, more specifically, to a method of repairing cracked aircraft structures utilizing a repair patch layup including a composite layer and a metal layer.
The necessity for repairing cracks in aircraft structures is well known. Aircraft, by nature are highly stressed from localized structural loading on areas such as the wings as well as from repetitive loading such as take-off and landing sequences. In order to be viable, aircraft structures must be lightweight yet capable of sustaining extreme forces. As a result of the forces applied, cracks often develop. In the case of military aircraft, use in conflict often gives rise to battle damage which, if left unrepaired, can quickly lead to undesirable crack propagation. Such crack propagation, if left unchecked, can result in the aircraft becoming unserviceable, clearly an unattractive result.
In the past, cracks, holes and the like have been repaired using a sheet patch, oftentimes aluminum. The patch is retained in place by a multiplicity of rivets. For example, U.S. Pat. No. 5,424,105 to Stewart discloses an aircraft patching method using a metallic patch held in place by rivets. While patching methods such as these are often quite useful, they work best in areas having low or no net stress loading. In areas wherein stresses are imparted, the crack underlying the patch will sometimes propagate in the repaired material or through the patch material, negating the repair.
Another repair method commonly used today is patch bonding. For example, metallic structures can be repaired by adhesively bonding boron-epoxy or carbon-epoxy laminated patches over the area to be repaired. The patch is bonded in place by means of a localized elevated temperature process. Boron-epoxy patches exhibit high stiffness and strength and are useful in many situations. But, this repair method is also not without shortcomings. For example, bonding a boron-epoxy patch to an aluminum wing structure gives rise to the creation of significant residual thermal stresses resulting from the elevated temperature cure cycle, as well as from the lower temperatures encountered during operation of the aircraft due to the different coefficients of thermal expansion of these materials. These thermal stresses dramatically reduce the effectiveness of the repair, and can again render the patch ineffective.
Some recent attempts have been made to utilize fiber-metal laminates originally developed as a fuselage material as repair patch materials. While generally successful in reducing crack propagation, these repairs are sensitive to stress concentration at the edges of the repair patch.
Therefore, while the repair methods of the prior art are somewhat successful, a need exists for an improved patching method. Such a method would effectively negate residual stress loading in an area by imparting a loading effect opposite to the stress loading in the area to be repaired.
Accordingly, it is a primary object of the present invention to provide a method of repairing cracked aircraft structures overcoming the limitations and disadvantages of the prior art.
Another object of the present invention is to provide a method of repairing cracked aircraft structures that provides a satisfactory permanent repair in areas subject to high stress loading.
Yet another object of the present invention is to provide a method of repairing cracked aircraft structures that is simple to complete.
Still another object of the present invention is to provide a method of repairing cracked aircraft structures that effectively imparts a loading effect opposite to the net loading on the area to be repaired.
It is still another object of the present invention to provide a method of repairing cracked aircraft structures that utilizes commonly used repair materials.
These and other objects of the invention will become apparent as the description of the representative embodiments proceeds.
In accordance with the foregoing principles and objects of the invention, a method of repairing cracked aircraft structures is described. The method includes determining the net loading of an area to be repaired and then identifying a repair patch layup that will impart a loading effect opposite thereto.
As is known, aircraft structures such as the fuselage and wings are highly stressed and are often subject to cracking. Military planes are perhaps even more susceptible to structure cracking than commercial aircraft due to the extreme usage military aircraft are subjected to and battle damage crack propagation. While the known methods of crack repair are quite useful in certain situations, they can themselves fail due to net loading on the repaired area. This net loading can be due to operational force loading on the aircraft structure from aircraft operation as well as residual thermal loading imparted by a difference in coefficients of thermal expansion between the patch material and the material of aircraft structure.
It is a well known fact that materials expand and contract with changes in temperature. A commonly used quantifier of this natural phenomenon is the coefficient of thermal expansion or (CTE). Generally, CTE can be thought of as, a ratio of the change in length of a line segment in a body per unit of temperature change to its length at a reference temperature. Dissimilar materials have different CTEs and the union of dissimilar materials such as is found in patching applications can impart a residual thermal loading effect between the materials as they expand and contract at different rates. This residual thermal loading effect can be quite undesirable for the repair.
In addition to the residual thermal loading effect imparted by the union of dissimilar patching materials, the operational loading of the structure to be repaired will often negatively impact the repair as well. This can be a significant problem resulting in patch disbonding or delamination or can even result in the crack propagating in the repaired material or through the patch material.
Advantageously, the method of the present invention takes into account the effect of the net loading of an area to be repaired. A repair patch layup consisting of a low CTE composite layer and a high CTE metal layer is identified such that the repair patch layup will impart a loading effect opposite thereto.